1. Technical Field
This invention applies to gas turbine rotor blades in general, and to cooled gas turbine rotor blades in particular.
2. Background Information
Turbine sections within an axial flow turbine engine include rotor assemblies that each includes a rotating disc and a number of rotor blades circumferentially disposed around the disk. Rotor blades include an airfoil portion for positioning within the gas path through the engine. Because the temperatures within the gas path very often negatively affect the durability of the airfoil, it is known to cool an airfoil by passing cooling air through the airfoil. The cooled air helps decrease the temperature of the airfoil material and thereby increase its durability.
Prior art cooled rotor blades very often utilize internal passage configurations that include a leading edge passage that either dead-ends adjacent the tip, or is connected to the tip by a cooling aperture, or is connected to an axially extending passage that dead-ends prior to the trailing edge. All of these internal passage configurations suffer from airflow stagnation regions, or regions of relatively low velocity flow that inhibit internal convective cooling. The airfoil wall regions adjacent these regions of low cooling effectiveness are typically at a higher temperature than other regions of the airfoil, and are therefore more prone to undesirable oxidation, thermal mechanical fatigue (TMF), creep, and erosion.
What is needed, therefore, is an airfoil having an internal passage configuration that promotes desirable cooling of the airfoil and thereby increases the durability of the blade.